Method for setting a gear ratio of a fan drive gear system of a gas turbine engine

ABSTRACT

A gas turbine engine includes a fan section including a fan rotatable about an axis. A speed reduction device is connected to the fan. The speed reduction device includes a planetary fan drive gear system with a planet gear ratio of at least 2.5. A bypass ratio is greater than about 11.0.

CROSS-REFERENCE TO RELATED APPLICATIONS

This disclosure is a continuation of U.S. patent application Ser. No.13/758,086 filed Feb. 4, 2013.

BACKGROUND

This disclosure relates to a gas turbine engine, and more particularlyto a method for setting a gear ratio of a fan drive gear system of a gasturbine engine.

A gas turbine engine may include a fan section, a compressor section, acombustor section, and a turbine section. Air entering the compressorsection is compressed and delivered into the combustor section where itis mixed with fuel and ignited to generate a high-speed exhaust gasflow. The high-speed exhaust gas flow expands through the turbinesection to drive the compressor and the fan section. Among othervariations, the compressor section can include low and high pressurecompressors, and the turbine section can include low and high pressureturbines.

Typically, a high pressure turbine drives a high pressure compressorthrough an outer shaft to form a high spool, and a low pressure turbinedrives a low pressure compressor through an inner shaft to form a lowspool. The fan section may also be driven by the inner shaft. A directdrive gas turbine engine may include a fan section driven by the lowspool such that a low pressure compressor, low pressure turbine, and fansection rotate at a common speed in a common direction.

A speed reduction device, which may be a fan drive gear system or othermechanism, may be utilized to drive the fan section such that the fansection may rotate at a speed different than the turbine section. Thisallows for an overall increase in propulsive efficiency of the engine.In such engine architectures, a shaft driven by one of the turbinesections provides an input to the speed reduction device that drives thefan section at a reduced speed such that both the turbine section andthe fan section can rotate at closer to optimal speeds.

Although gas turbine engines utilizing speed change mechanisms aregenerally known to be capable of improved propulsive efficiency relativeto conventional engines, gas turbine engine manufacturers continue toseek further improvements to engine performance including improvementsto thermal, transfer and propulsive efficiencies.

SUMMARY

In one exemplary embodiment, a gas turbine engine includes a fan sectionincluding a fan rotatable about an axis. A speed reduction device isconnected to the fan. The speed reduction device includes a planetaryfan drive gear system with a planet gear ratio of at least 2.5. A bypassratio is greater than about 11.0.

In a further embodiment of the above, the gear ratio is less than orequal to 5.0.

In a further embodiment of any of the above, the gas turbine engineincludes a fan pressure ratio that is below 1.7.

In a further embodiment of any of the above, the gas turbine engineincludes a fan pressure ratio that is below 1.48.

In a further embodiment of any of the above, the fan blade tip speed ofthe fan section is greater than about 1000 ft/sec and less than about1400 ft/sec.

In a further embodiment of any of the above, the planetary fan drivegear system includes a sun gear, a plurality of planetary gears, a ringgear, and a carrier.

In a further embodiment of any of the above, each of the plurality ofplanetary gears includes at least one bearing.

In a further embodiment of any of the above, the ring gear is fixed fromrotation.

In a further embodiment of any of the above, a low pressure turbine ismechanically attached to the sun gear.

In a further embodiment of any of the above, a fan section ismechanically attached to the carrier.

In a further embodiment of any of the above, an input of the speedreduction device is rotatable in a first direction and an output of thespeed reduction device is rotatable in the same first direction.

In a further embodiment of any of the above, a low pressure turbinesection is in communication with the speed reduction device. The lowpressure turbine section includes at least three stages.

In a further embodiment of any of the above, the bypass ratio is lessthan about 22.0.

In another exemplary embodiment, a method of improving performance of agas turbine engine includes determining fan tip speed boundaryconditions for at least one fan blade of a fan section, determiningrotor boundary conditions for a rotor of a low pressure turbine, andutilizing stress level constraints in the rotor of the low pressureturbine and the at least one fan blade to determine if the rotary speedof the fan section and the low pressure turbine will meet a desirednumber of operating cycles. A bypass ratio is greater than about 6.0. Aspeed reduction device connects the fan section and the low pressureturbine and includes a planetary gear ratio of at least about 2.5.

In a further embodiment of any of the above, the planetary gear ratio isless than about 5.0.

In a further embodiment of any of the above, a fan pressure ratio isbelow 1.7.

In a further embodiment of any of the above, a fan pressure ratio isbelow 1.48.

In a further embodiment of any of the above, the bypass ratio is greaterthan about 11 and less than about 22.

In a further embodiment of any of the above, a fan blade tip speed of atleast one fan blade is less than 1400 fps.

In a further embodiment of any of the above, if a stress level in therotor or at least one fan blade is too high to meet a desired number ofoperating cycles, a gear ratio of a gear reduction device is lowered andthe number of stages of the low pressure turbine is increased.

In a further embodiment of any of the above, if a stress level in therotor or at least one fan blade is too high to meet a desired number ofoperating cycles, a gear ratio of a gear reduction device is lowered andan annular area of the low pressure turbine is increased.

In another exemplary embodiment, a fan drive gear module for a gasturbine engine includes a planetary fan drive gear system with a speedreduction ratio of at least 2.5. The speed reduction device isconfigured to drive a fan section with a bypass ratio greater than about11.0.

In a further embodiment of the above, the speed reduction ratio is lessthan or equal to 5.0.

In a further embodiment of any of the above, the planetary fan drivegear system is configured to drive a fan section with a fan blade tipspeed greater than about 1000 ft/sec and less than about 1400 ft/sec.

In a further embodiment of any of the above, the bypass ratio is lessthan about 22.0.

In another exemplary embodiment, a method of designing a gas turbineengine includes selecting fan tip speed boundary conditions for at leastone fan blade of a fan section of a gas turbine engine, selecting rotorboundary conditions for a rotor of a fan drive turbine of the gasturbine engine and determining stress level constraints in the rotor ofthe fan drive turbine and the at least one fan blade to determine if therotary speed of the fan section and the fan drive turbine will meet adesired number of operating cycles wherein a bypass ratio is greaterthan about 6.0 and connecting the fan section and the fan drive turbinethrough a speed reduction device that includes a planetary gear ratio ofat least about 2.5.

In a further embodiment of the above, the planetary gear ratio is lessthan about 5.0.

In a further embodiment of any of the above, the bypass ratio of the gasturbine engine is greater than about 11 and less than about 22.

In a further embodiment of any of the above, the method includeslowering the speed reduction ratio of the speed reduction device andincreasing a number of stages of the fan drive turbine responsive todetermining that a stress level in the rotor or the at least one fanblade is outside a predefined number of operating cycles.

In a further embodiment of any of the above, the method includeslowering the speed reduction ratio of the speed reduction device andincreasing an annular area of the fan drive turbine responsive todetermining that a stress level in the rotor or the at least one fanblade is outside a predefined number of operating cycles.

The various features and advantages of this disclosure will becomeapparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a schematic, cross-sectional view of an example gasturbine engine.

FIG. 2 illustrates a schematic view of one configuration of a low speedspool that can be incorporated into a gas turbine engine.

FIG. 3 illustrates a fan drive gear system that can be incorporated intoa gas turbine engine.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The exemplarygas turbine engine 20 is a two-spool turbofan engine that generallyincorporates a fan section 22, a compressor section 24, a combustorsection 26 and a turbine section 28. Alternative engines might includean augmentor section (not shown) among other systems or features. Thefan section 22 drives air along a bypass flow path B, while thecompressor section 24 drives air along a core flow path C forcompression and communication into the combustor section 26. The hotcombustion gases generated in the combustor section 26 are expandedthrough the turbine section 28. Although depicted as a two-spoolturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto two-spool turbofan engines and these teachings could extend to othertypes of engines, including but not limited to, three-spool enginearchitectures.

The exemplary gas turbine engine 20 generally includes a low speed spool30 and a high speed spool 32 mounted for rotation about an enginecenterline longitudinal axis A. The low speed spool 30 and the highspeed spool 32 may be mounted relative to an engine static structure 33via several bearing systems 31. It should be understood that otherbearing systems 31 may alternatively or additionally be provided, andthe location of bearing systems 31 may be varied as appropriate to theapplication.

The low speed spool 30 generally includes an inner shaft 34 thatinterconnects a fan 36, a low pressure compressor 38 and a low pressureturbine 39. The inner shaft 34 can be connected to the fan 36 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 45, such as a fan drive gear system50 (see FIGS. 2 and 3). The speed change mechanism drives the fan 36 ata lower speed than the low speed spool 30. The high speed spool 32includes an outer shaft 35 that interconnects a high pressure compressor37 and a high pressure turbine 40. In this embodiment, the inner shaft34 and the outer shaft 35 are supported at various axial locations bybearing systems 31 positioned within the engine static structure 33.

A combustor 42 is arranged in exemplary gas turbine 20 between the highpressure compressor 37 and the high pressure turbine 40. A mid-turbineframe 44 may be arranged generally between the high pressure turbine 40and the low pressure turbine 39. The mid-turbine frame 44 can supportone or more bearing systems 31 of the turbine section 28. Themid-turbine frame 44 may include one or more airfoils 46 that extendwithin the core flow path C. It will be appreciated that each of thepositions of the fan section 22, compressor section 24, combustorsection 26, turbine section 28, and fan drive gear system 50 may bevaried. For example, gear system 50 may be located aft of combustorsection 26 or even aft of turbine section 28, and fan section 22 may bepositioned forward or aft of the location of gear system 50.

The inner shaft 34 and the outer shaft 35 are concentric and rotate viathe bearing systems 31 about the engine centerline longitudinal axis A,which is co-linear with their longitudinal axes. The core airflow iscompressed by the low pressure compressor 38 and the high pressurecompressor 37, is mixed with fuel and burned in the combustor 42, and isthen expanded over the high pressure turbine 40 and the low pressureturbine 39. The high pressure turbine 40 and the low pressure turbine 39rotationally drive the respective high speed spool 32 and the low speedspool 30 in response to the expansion.

In a non-limiting embodiment, the gas turbine engine 20 is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20bypass ratio is greater than about six (6:1). The geared architecture 45can include an epicyclic gear train, such as a planetary gear system, astar gear system, or other gear system. The geared architecture 45enables operation of the low speed spool 30 at higher speeds, which canenable an increase in the operational efficiency of the low pressurecompressor 38 and low pressure turbine 39, and render increased pressurein a fewer number of stages.

The pressure ratio of the low pressure turbine 39 can be pressuremeasured prior to the inlet of the low pressure turbine 39 as related tothe pressure at the outlet of the low pressure turbine 39 and prior toan exhaust nozzle of the gas turbine engine 20. In one non-limitingembodiment, the bypass ratio of the gas turbine engine 20 is greaterthan about ten (10:1), the fan diameter is significantly larger thanthat of the low pressure compressor 38, and the low pressure turbine 39has a pressure ratio that is greater than about five (5:1). In anothernon-limiting embodiment, the bypass ratio is greater than 11 and lessthan 22, or greater than 13 and less than 20. It should be understood,however, that the above parameters are only exemplary of a gearedarchitecture engine or other engine using a speed change mechanism, andthat the present disclosure is applicable to other gas turbine engines,including direct drive turbofans. In one non-limiting embodiment, thelow pressure turbine 39 includes at least one stage and no more thaneight stages, or at least three stages and no more than six stages. Inanother non-limiting embodiment, the low pressure turbine 39 includes atleast three stages and no more than four stages.

In this embodiment of the exemplary gas turbine engine 20, a significantamount of thrust is provided by the bypass flow path B due to the highbypass ratio. The fan section 22 of the gas turbine engine 20 isdesigned for a particular flight condition—typically cruise at about 0.8Mach and about 35,000 feet. This flight condition, with the gas turbineengine 20 at its best fuel consumption, is also known as bucket cruiseThrust Specific Fuel Consumption (TSFC). TSFC is an industry standardparameter of fuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. The low FanPressure Ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45. In another non-limitingembodiment of the example gas turbine engine 20, the Fan Pressure Ratiois less than 1.38 and greater than 1.25. In another non-limitingembodiment, the fan pressure ratio is less than 1.48. In anothernon-limiting embodiment, the fan pressure ratio is less than 1.52. Inanother non-limiting embodiment, the fan pressure ratio is less than1.7. Low Corrected Fan Tip Speed is the actual fan tip speed divided byan industry standard temperature correction of [(Tram ° R)/(518.7°R)]^(0.5), where T represents the ambient temperature in degreesRankine. The Low Corrected Fan Tip Speed according to one non-limitingembodiment of the example gas turbine engine 20 is less than about 1150fps (351 m/s). The Low Corrected Fan Tip Speed according to anothernon-limiting embodiment of the example gas turbine engine 20 is lessthan about 1400 fps (427 m/s). The Low Corrected Fan Tip Speed accordingto another non-limiting embodiment of the example gas turbine engine 20is greater than about 1000 fps (305 m/s).

FIG. 2 schematically illustrates the low speed spool 30 of the gasturbine engine 20. The low speed spool 30 includes the fan 36, the lowpressure compressor 38, and the low pressure turbine 39. The inner shaft34 interconnects the fan 36, the low pressure compressor 38, and the lowpressure turbine 39. The inner shaft 34 is connected to the fan 36through the fan drive gear system 50. In this embodiment, the fan drivegear system 50 provides for co-rotation of the low pressure turbine 39and the fan 36. For example, the fan 36 rotates in a first direction D1and the low pressure turbine 39 rotates in the same first direction D1as the fan 36.

FIG. 3 illustrates one example embodiment of the fan drive gear system50 incorporated into the gas turbine engine 20 to provide forco-rotation of the fan 36 and the low pressure turbine 39. In thisembodiment, the fan drive gear system 50 includes a planetary gearsystem having a sun gear 52, a fixed ring gear 54 disposed about the sungear 52, and a plurality of planetary gears 56 having journal bearings57 positioned between the sun gear 52 and the ring gear 54. A carrier 58carries and is attached to each of the planetary gears 56. In thisembodiment, the fixed ring gear 54 does not rotate and is connected to agrounded structure 55 of the gas turbine engine 20.

The sun gear 52 receives an input from the low pressure turbine 39 (seeFIG. 2) and rotates in a first direction D1 thereby turning theplurality of planetary gears 56 in a second direction D2 that isopposite of the first direction D1. Movement of the plurality ofplanetary gears 56 is transmitted to the carrier 58, which rotates inthe first direction D1. The carrier 58 is connected to the fan 36 forrotating the fan 36 (see FIG. 2) in the first direction D1.

A planet system gear ratio of the fan drive gear system 50 is determinedby measuring a diameter of the ring gear 54 and dividing that diameterby a diameter of the sun gear 52 and adding one to the quotient. In oneembodiment, the planet system gear ratio of the fan drive gear system 50is between 2.5 and 5.0. When the planetary system gear ratio is below2.5, the sun gear 52 is relatively much larger than the planetary gears56. This size differential reduces the load the planetary gears 56 arecapable of carrying because of the reduction in size of the journalbearings 57. When the system gear ratio is above 5.0, the sun gear 52 isrelatively much smaller than the planetary gears 56. This sizedifferential increases the size of the planetary gear 56 journalbearings 57 but reduces the load the sun gear 52 is capable of carryingbecause of its reduced size and number of teeth. Alternatively, rollerbearings could be used in place of journal bearings 57.

Improving performance of the gas turbine engine 20 begins by determiningfan tip speed boundary conditions for at least one fan blade of the fan36 to define the speed of the tip of the fan blade. The maximum fandiameter is determined based on the projected fuel burn derived frombalancing engine efficiency, mass of air through the bypass flow path B,and engine weight increase due to the size of the fan blades.

Boundary conditions are then determined for the rotor of each stage ofthe low pressure turbine 39 to define the speed of the rotor tip and todefine the size of the rotor and the number of stages in the lowpressure turbine 39 based on the efficiency of low pressure turbine 39and the low pressure compressor 38.

Constraints regarding stress levels in the rotor and the fan blade areutilized to determine if the rotary speed of the fan 36 and the lowpressure turbine 39 will meet a desired number of operating life cycles.If the stress levels in the rotor or the fan blade are too high, thegear ratio of the fan drive gear system 50 can be lowered and the numberof stages of the low pressure turbine 39 or annular area of the lowpressure turbine 39 can be increased.

Although the different non-limiting embodiments are illustrated ashaving specific components, the embodiments of this disclosure are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed and illustrated in these exemplary embodiments,other arrangements could also benefit from the teachings of thisdisclosure.

The foregoing description shall be interpreted as illustrative and notin any limiting sense. A worker of ordinary skill in the art wouldunderstand that certain modifications could come within the scope ofthis disclosure. For these reasons, the following claim should bestudied to determine the true scope and content of this disclosure.

What is claimed is:
 1. A gas turbine engine comprising: a fan sectionincluding a fan rotatable about an axis; and a speed reduction deviceconnected to the fan, wherein the speed reduction device includes aplanetary fan drive gear system with a planet gear ratio of at least2.5, wherein a bypass ratio is greater than about 11.0.
 2. The gasturbine engine of claim 1, wherein the gear ratio is less than or equalto 5.0.
 3. The gas turbine engine of claim 2, including a fan pressureratio that is below 1.7.
 4. The gas turbine engine of claim 2, includinga fan pressure ratio that is below 1.48.
 5. The gas turbine engine ofclaim 4, wherein the fan blade tip speed of the fan section is greaterthan about 1000 ft/sec and less than about 1400 ft/sec.
 6. The gasturbine engine of claim 1, wherein the planetary fan drive gear systemincludes a sun gear, a plurality of planetary gears, a ring gear, and acarrier.
 7. The gas turbine engine of claim 6, wherein each of theplurality of planetary gears includes at least one bearing.
 8. The gasturbine engine of claim 7, wherein the ring gear is fixed from rotation.9. The gas turbine engine of claim 8, wherein a low pressure turbine ismechanically attached to the sun gear.
 10. The gas turbine engine ofclaim 9, wherein a fan section is mechanically attached to the carrier.11. The gas turbine engine of claim 10, wherein an input of the speedreduction device is rotatable in a first direction and an output of thespeed reduction device is rotatable in the same first direction.
 12. Thegas turbine engine of claim 11, including a low pressure turbine sectionin communication with the speed reduction device, wherein the lowpressure turbine section includes at least three stages.
 13. The gasturbine engine of claim 1, wherein the bypass ratio is less than about22.0.
 14. A method of improving performance of a gas turbine enginecomprising: determining fan tip speed boundary conditions for at leastone fan blade of a fan section; determining rotor boundary conditionsfor a rotor of a low pressure turbine; and utilizing stress levelconstraints in the rotor of the low pressure turbine and the at leastone fan blade to determine if the rotary speed of the fan section andthe low pressure turbine will meet a desired number of operating cycles,wherein a bypass ratio is greater than about 6.0, wherein a speedreduction device connects the fan section and the low pressure turbineand includes a planetary gear ratio of at least about 2.5.
 15. Themethod of claim 14, wherein the planetary gear ratio is less than about5.0.
 16. The method of claim 15, wherein a fan pressure ratio is below1.7.
 17. The method of claim 16, wherein a fan pressure ratio is below1.48.
 18. The method of claim 17, wherein the bypass ratio is greaterthan about 11 and less than about
 22. 19. The method of claim 18,wherein a fan blade tip speed of the at least one fan blade is less than1400 fps.
 20. The method of claim 14, wherein if a stress level in therotor or the at least one fan blade is too high to meet a desired numberof operating cycles, a gear ratio of a gear reduction device is loweredand the number of stages of the low pressure turbine is increased. 21.The method of claim 20, wherein if a stress level in the rotor or the atleast one fan blade is too high to meet a desired number of operatingcycles, a gear ratio of a gear reduction device is lowered and anannular area of the low pressure turbine is increased.
 22. A fan drivegear module for a gas turbine engine comprising: a planetary fan drivegear system with a speed reduction ratio of at least 2.5, wherein thespeed reduction device is configured to drive a fan section with abypass ratio greater than about 11.0.
 23. The fan drive gear module ofclaim 22, wherein the speed reduction ratio is less than or equal to5.0.
 24. The fan drive gear module of claim 22, wherein the planetaryfan drive gear system is configured to drive a fan section with a fanblade tip speed greater than about 1000 ft/sec and less than about 1400ft/sec.
 25. The fan drive gear module of claim 22, wherein the bypassratio is less than about 22.0.
 26. A method of designing a gas turbineengine comprising: selecting fan tip speed boundary conditions for atleast one fan blade of a fan section of a gas turbine engine; selectingrotor boundary conditions for a rotor of a fan drive turbine of the gasturbine engine; determining stress level constraints in the rotor of thefan drive turbine and the at least one fan blade to determine if therotary speed of the fan section and the fan drive turbine will meet adesired number of operating cycles, wherein a bypass ratio is greaterthan about 6.0; and connecting the fan section and the fan drive turbinethrough a speed reduction device that includes a planetary gear ratio ofat least about 2.5.
 27. The method of claim 26, wherein the planetarygear ratio is less than about 5.0.
 28. The method of claim 26, whereinthe bypass ratio of the gas turbine engine is greater than about 11 andless than about
 22. 29. The method of claim 26, including lowering thespeed reduction ratio of the speed reduction device and increasing anumber of stages of the fan drive turbine responsive to determining thata stress level in the rotor or the at least one fan blade is outside apredefined number of operating cycles.
 30. The method of claim 26,including lowering the speed reduction ratio of the speed reductiondevice and increasing an annular area of the fan drive turbineresponsive to determining that a stress level in the rotor or the atleast one fan blade is outside a predefined number of operating cycles.